Field of the Invention
The present invention relates to an axial flow compressor, a gas turbine including the same, and a stator blade of an axial flow compressor.
Background Art
In axial flow compressors, a rotor blade row and a stator blade row are formed of multiple rotor blades and multiple stator blades which are arranged in a circumferential direction of an annular channel through which a working fluid flows, and one stage consists of one set of a rotor blade row and a stator blade row. The axial flow compressors include multiple stages.
In recent years, the axial flow compressors have needed higher loading which compatibly satisfies a higher pressure ratio and cost saving achieved by reducing the number of stages. In a subsonic airfoil of a high loaded compressor, secondary flow increases due to a developed boundary layer on a wall surface (endwall on one end side of a blade row) on an inner peripheral side or an outer peripheral side of an annular channel where the blade is located. Consequently, pressure loss may increase due to flow stall (corner stall) in a corner portion formed between a blade surface and the wall surface of the channel. Therefore, in order to develop a high performance and high loaded compressor, it is an important task to create a high performance airfoil and channel wall surface contour capable of restraining the corner stall.
For example, as a stator blade of a compressor which can improve both efficiency and a stall margin of the compressor at the same time while flow separation is avoided in the vicinity of a channel wall surface (endwall on one end side of a blade row), JP-A-2001-132696 discloses a technology in which a chord length of a radial span central portion (waist) of a stator blade is set to be shorter than that of a blade tip or a blade hub, and in which a trailing edge of the blade is bowed.